Cooled turbine blade with inner spar

ABSTRACT

A cooled turbine blade having a base and an airfoil, the base including cooling air inlet and an internal cooling air passageway, and the airfoil including an internal heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a “skin” that encompasses a tip wall, an inner spar, a plurality of inner spar cooling fins extending from the inner spar to the skin, and a plurality of trailing edge cooling fins extending from the pressure side of the skin to the lift side of the skin aft of the inner spar.

TECHNICAL FIELD

The present disclosure generally pertains to gas turbine engines, and is more particularly directed toward a cooled turbine blade.

BACKGROUND

High performance gas turbine engines typically rely on increasing turbine inlet temperatures to increase both fuel economy and overall power ratings. These higher temperatures, if not compensated for, oxidize engine components and decrease component life. Component life has been increased by a number of techniques. Said techniques include internal cooling with air bled from an engine compressor section. Bleeding air results in efficiency loss however. In addition, stationary gas turbine engines typically may have less available compressed air than moving gas turbine engines.

U.S. Pat. No. 7,690,894 issued to Liang on Apr. 6, 2010 shows a ceramic core assembly for a serpentine flow circuit in a turbine blade. In particular, the disclosure of Liang is directed toward a turbine blade for use in a gas turbine engine having an internal serpentine flow cooling circuit with pin fins and trip strips to promote heat transfer for obtaining a thermally balanced blade sectional temperature distribution. The turbine blade is cooled by a 7-pass serpentine flow cooling circuit that extends from the leading edge and along the pressure side wall of the airfoil, into the trailing edge and then flows along the suction side wall ending just downstream from the leading edge where the 7-pass serpentine flow circuit started. Leading edge film cooling holes are supplied from the first leg of the serpentine while a row of trailing edge exit holes is supplied from the third leg which extends across both walls of the airfoil in the trailing edge.

The present disclosure is directed toward overcoming one or more of the problems discovered by the inventors.

SUMMARY OF THE DISCLOSURE

A cooled turbine blade is disclosed herein. The cooled turbine blade having a base and an airfoil, the base including cooling air inlet and an internal cooling air passageway, and the airfoil including an internal heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a skin that encompasses a tip wall, an inner spar, a plurality of inner spar cooling fins extending from the inner spar to the skin, a plurality of trailing edge cooling fins extending from the pressure side of the skin to the lift side of the skin aft of the inner spar, and a leading edge chamber. According to one embodiment, a cooled turbine blade, similar to the above but wherein the inner spar cooling fins are substantially parallel to each other along the mean camber line of the airfoil, is also disclosed herein. According to another embodiment, a cooled turbine blade, similar to the above but wherein the cooling air follows a single bend heat exchange path, is also disclosed herein.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration of an exemplary gas turbine engine.

FIG. 2 is an axial view of an exemplary turbine rotor assembly.

FIG. 3 is an isometric view of an exemplary turbine blade.

FIG. 4 is a cutaway side view of an exemplary turbine blade.

FIG. 5 is a sectional top view of an exemplary turbine blade, as taken along line 5-5 of FIG. 4.

DETAILED DESCRIPTION

FIG. 1 is a schematic illustration of an exemplary gas turbine engine. Some of the surfaces have been left out or exaggerated (here and in other figures) for clarity and ease of explanation. Also, the disclosure may reference a forward and an aft direction. Generally, all references to “forward” and “aft” are associated with the flow direction of primary air (i.e., air used in the combustion process), unless specified otherwise. For example, forward is “upstream” relative to primary air flow, and aft is “downstream” relative to primary air flow.

In addition, the disclosure may generally reference a center axis 95 of rotation of the gas turbine engine, which may be generally defined by the longitudinal axis of its shaft 120 (supported by a plurality of bearing assemblies 150). The center axis 95 may be common to or shared with various other engine concentric components. All references to radial, axial, and circumferential directions and measures refer to center axis 95, unless specified otherwise, and terms such as “inner” and “outer” generally indicate a lesser or greater radial distance from, wherein a radial 96 may be in any direction perpendicular and radiating outward from center axis 95.

Structurally, a gas turbine engine 100 includes an inlet 110, a gas producer or “compressor” 200, a combustor 300, a turbine 400, an exhaust 500, and a power output coupling 600. The compressor 200 includes one or more compressor rotor assemblies 220. The combustor 300 includes one or more injectors 350 and includes one or more combustion chambers 390. The turbine 400 includes one or more turbine rotor assemblies 420. The exhaust includes an exhaust diffuser 520 and an exhaust collector 550.

As illustrated, both compressor rotor assembly 220 and turbine rotor assembly 420 are axial flow rotor assemblies, where each rotor assembly includes a rotor disk that is circumferentially populated with a plurality of airfoils (“rotor blades”). When installed, the rotor blades associated with one rotor disk are axially separated from the rotor blades associated with an adjacent disk by stationary vanes (“stator vanes” or “stators”) 250, 450 circumferentially distributed in an annular casing.

Functionally, a gas (typically air 10) enters the inlet 110 as a “working fluid”, and is compressed by the compressor 200. In the compressor 200, the working fluid is compressed in an annular flow path 115 by the series of compressor rotor assemblies 220. In particular, the air 10 is compressed in numbered “stages”, the stages being associated with each compressor rotor assembly 220. For example, “4th stage air” may be associated with the 4th compressor rotor assembly 220 in the downstream or “aft” direction—going from the inlet 110 towards the exhaust 500). Likewise, each turbine rotor assembly 420 may be associated with a numbered stage. For example, first stage turbine rotor assembly 421 is the forward most of the turbine rotor assemblies 420. However, other numbering/naming conventions may also be used.

Once compressed air 10 leaves the compressor 200, it enters the combustor 300, where it is diffused and fuel 20 is added. Air 10 and fuel 20 are injected into the combustion chamber 390 via injector 350 and ignited. After the combustion reaction, energy is then extracted from the combusted fuel/air mixture via the turbine 400 by each stage of the series of turbine rotor assemblies 420. Exhaust gas 90 may then be diffused in exhaust diffuser 520 and collected, redirected, and exit the system via an exhaust collector 550. Exhaust gas 90 may also be further processed (e.g., to reduce harmful emissions, and/or to recover heat from the exhaust gas 90).

One or more of the above components (or their subcomponents) may be made from stainless steel and/or durable, high temperature materials known as “superalloys”. A superalloy, or high-performance alloy, is an alloy that exhibits excellent mechanical strength and creep resistance at high temperatures, good surface stability, and corrosion and oxidation resistance. Superalloys may include materials such as HASTELLOY, INCONEL, WASPALOY, RENE alloys, HAYNES alloys, INCOLOY, MP98T, TMS alloys, and CMSX single crystal alloys.

FIG. 2 is an axial view of an exemplary turbine rotor assembly. In particular, first stage turbine rotor assembly 421 schematically illustrated in FIG. 1 is shown here in greater detail, but in isolation from the rest of gas turbine engine 100. First stage turbine rotor assembly 421 includes a turbine rotor disk 430 that is circumferentially populated with a plurality of turbine blades configured to receive cooling air (“cooled turbine blades” 440) and a plurality of dampers 426. Here, for illustration purposes, turbine rotor disk 430 is shown depopulated of all but three cooled turbine blades 440 and three dampers 426.

Each cooled turbine blade 440 may include a base 442 including a platform 443 and a blade root 480. For example, the blade root 480 may incorporate “fir tree”, “bulb”, or “dove tail” roots, to list a few. Correspondingly, the turbine rotor disk 430 may include a plurality of circumferentially distributed slots or “blade attachment grooves” 432 configured to receive and retain each cooled turbine blade 440. In particular, the blade attachment grooves 432 may be configured to mate with the blade root 480, both having a reciprocal shape with each other. In addition the blade attachment grooves 432 may be slideably engaged with the blade attachment grooves 432, for example, in a forward-to-aft direction.

Being proximate the combustor 300 (FIG. 1), the first stage turbine rotor assembly 421 may incorporate active cooling. In particular, compressed cooling air may be internally supplied to each cooled turbine blade 440 as well as predetermined portions of the turbine rotor disk 430. For example, here turbine rotor disk 430 engages the cooled turbine blade 440 such that a cooling air cavity 433 is formed between the blade attachment grooves 432 and the blade root 480. In other embodiments, other stages of the turbine may incorporate active cooling as well.

When a pair of cooled turbine blades 440 is mounted in adjacent blade attachment grooves 432 of turbine rotor disk 430, an under-platform cavity may be formed above the circumferential outer edge of turbine rotor disk 430, between shanks of adjacent blade roots 480, and below their adjacent platforms 443, respectively. As such, each damper 426 may be configured to fit this under-platform cavity. Alternately, where the platforms are flush with circumferential outer edge of turbine rotor disk 430, and/or the under-platform cavity is sufficiently small, the damper 426 may be omitted entirely.

Here, as illustrated, each damper 426 may be configured to constrain received cooling air such that a positive pressure may be created within under-platform cavity to suppress the ingress of hot gases from the turbine. Additionally, damper 426 may be further configured to regulate the flow of cooling air to components downstream of the first stage turbine rotor assembly 421. For example, damper 426 may include one or more aft plate apertures in its aft face. Certain features of the illustration may be simplified and/or differ from a production part for clarity.

Each damper 426 may be configured to be assembled with the turbine rotor disk 430 during assembly of first stage turbine rotor assembly 421, for example, by a press fit. In addition, the damper 426 may form at least a partial seal with the adjacent cooled turbine blades 440. Furthermore, one or more axial faces of damper 426 may be sized to provide sufficient clearance to permit each cooled turbine blade 440 to slide into the blade attachment grooves 432, past the damper 426 without interference after installation of the damper 426.

FIG. 3 is an isometric view of the turbine blade of FIG. 2. As described above, the cooled turbine blade 440 may include a base 442 having a platform 443 and a blade root 480. Each cooled turbine blade 440 may further include an airfoil 441 extending radially outward from the platform 443. The airfoil 441 may have a complex, geometry that varies radially. For example the cross section of the airfoil 441 may lengthen, thicken, twist, and/or change shape as it radially approaches the platform 443 inward from the tip end 445. The overall shape of airfoil 441 may also vary from application to application.

The cooled turbine blade 440 is generally described herein with reference to its installation and operation. In particular, the cooled turbine blade 440 is described with reference to both a radial 96 of center axis 95 (FIG. 1) and the aerodynamic features of the airfoil 441. The aerodynamic features of the airfoil 441 include a leading edge 446, a trailing edge 447, a pressure side 448, a lift side 449, and its mean camber line 474. The mean camber line 474 is generally defined as the line running along the center of the airfoil from the leading edge 446 to the trailing edge 447. It can be thought of as the average of the pressure side 448 and lift side 449 of the airfoil shape. As discussed above, airfoil 474 also extends radially between the platform 443 and the tip end 445. Accordingly, the mean camber line 474 herein includes the entire camber sheet continuing from the platform 443 to the tip end 445.

Accordingly, when describing the cooled turbine blade 440 as a unit, the inward direction is generally radially inward toward the center axis 95 (FIG. 1), with its associated end called the “root end” 444. Likewise is the outward direction is generally radially outward from the center axis 95 (FIG. 1), with its associated end called the “tip end” 445. When describing the platform 443, the forward edge 484 and the aft edge 485 of the platform 443 are associated the forward and aft axial directions of the center axis 95 (FIG. 1), as described above.

In addition, when describing the airfoil 441, the forward and aft directions are generally measured between its leading edge 446 (forward) and its trailing edge 447 (aft), along the mean camber line 474 (artificially treating the mean camber line 474 as linear). When describing the flow features of the airfoil 441, the inward and outward directions are generally measured in the radial direction relative to the center axis 95 (FIG. 1). However, when describing the thermodynamic features of the airfoil 441 (particularly those associated with the inner spar 462 (FIG. 5)), the inward and outward directions are generally measured in a plane perpendicular to a radial 96 of center axis 95 (FIG. 1) with “inward” being toward the mean camber line 474 and “outward” being toward the “skin” 460 of the airfoil 441.

Finally, certain traditional aerodynamics terms may be used from time to time herein for clarity, but without being limiting. For example, while it will be discussed that the airfoil 441 (along with the entire cooled turbine blade 440) may be made as a single metal casting, the outer surface of the airfoil 441 (along with its thickness) is descriptively called herein the “skin” 460 of the airfoil 441.

FIG. 4 is a cutaway side view of the turbine blade of FIG. 3. In particular, the cooled turbine blade 440 of FIG. 3 is shown here with sections of the skin 460 removed from the pressure side 448 of the airfoil 441, exposing its internal structure and cooling paths. For example, the airfoil 441 may include a composite flow path made up of multiple subdivisions and cooling structures. Similarly, a section of the base 442 has been removed to expose portions of a cooling air passageway 482, internal to the base 442.

As described above, the cooled turbine blade 440 may include an airfoil 441 and a base 442. The base 442 may include the platform 443, the blade root 480, and one or more cooling air inlet(s) 481. The airfoil 441 interfaces with the base 442 and may include the skin 460, a tip wall 461, and the cooling air outlet 471.

Compressed secondary air may be routed into one or more cooling air inlet(s) 481 in the base 442 of cooled turbine blade 440 as cooling air 15. The one or more cooling air inlet(s) 481 may be at any convenient location. For example, here the cooling air inlet 481 is located in the blade root 480. Alternately, cooling air 15 may be received in a shank area radially outward from the blade root 480 but radially inward from the platform 443.

Within the base 442, the cooled turbine blade 440 include the cooling air passageway 482 that is configured to route cooling air 15 from the one or more cooling air inlet(s) 481, through the base, and into the airfoil 441. The cooling air passageway 482 may be configured to translate the cooling air 15 in two dimensions (i.e., not merely in the plane of the figure) as it travels radially up (i.e., generally in the direction of a radial 96 of the center axis 95 (FIG. 1)) towards the airfoil 441. Moreover, the cooling air passageway 482 may be structured to receive the cooling air 15 from a generally rectilinear cooling air inlet 481 and smoothly “reshape” it fit the curvature and shape of the airfoil 441. In addition, the cooling air passageway 482 may be subdivided into a plurality of subpassages. As illustrated, the subdivisions may be evenly spaced, for example.

Within the skin 460 of the airfoil 441, several internal structures are viewable. In particular, airfoil 441 may include a tip wall 461, an inner spar 462, a leading edge chamber 463, one or more section divider(s) 464, one or more rib(s) 465, one or more air deflector(s) 466, and a plurality of inner spar cooling fins 467. In addition, airfoil 441 may include a perforated trailing edge rib 468 and a plurality of trailing edge cooling fins 469. Together with the skin 460, these structures may form a single-bend heat exchange path 470 within the airfoil 441.

The internal structures making up the single-bend heat exchange path 470 may subdivide the single-bend heat exchange path 470 into multiple discrete sub-passageways or “sections”. For example, although single-bend heat exchange path 470 is shown by a representative path of cooling air 15, three completely separated sections are illustrated (i.e., separated by section dividers 464) here on the pressure side 448 of cooled turbine blade 440. Furthermore, in the particular embodiment illustrated, a total of six sub-passageways (including leading edge chamber 463) are identifiable.

With regard to the airfoil structures, the tip wall 461 extends across the airfoil 441 and may be configured to redirect cooling air 15 from escaping through the tip end 445. In addition, one embodiment of the tip end 445 is the tip wall 461. Moreover, tip end 445 may be formed as a shared structure, such as a joining of the pressure side 448 and the lift side 449 of the airfoil 441. According to one embodiment, the tip wall 461 may be recessed inward such that it is not flush with the tip of the airfoil 441. According to one embodiment, the tip wall 461 may include one or more perforations (not shown) such that a small quantity of the cooling air 15 may be bled off for film cooling of the tip end 445.

The inner spar 462 may extend from the base 442 radially outward to the tip wall 461, between the pressure side 448 (FIG. 3) and the lift side 449 (FIG. 3) of the skin 460. In addition, the inner spar 462 may extend between the leading edge 446 and the trailing edge 447, parallel with, and generally following, the mean camber line 474 (FIG. 3) of the airfoil 441, and terminating with inner spar trailing edge 476. Accordingly, the inner spar 462 may be configured to bifurcate a portion or all of the airfoil 441 generally along its mean camber line 474 (FIG. 3) and between the pressure side 448 and the lift side 449. Also, the inner spar 462 may be solid (non-perforated) or substantially solid, such that cooling air 15 cannot pass.

According to one embodiment, the inner spar 462 may extend less than the entire length of the mean camber line 474. In particular the inner spar 462 may extend less than ninety percent of the mean camber line 474 and may exclude the leading edge chamber 463 entirely. For example, the inner spar 462 may extend from the leading edge chamber 463, downstream to the plurality of trailing edge cooling fins 469. In addition, the inner spar 462 may have a length within the range of seventy to eighty percent, or approximately three quarters the length of, and along, the mean camber line 474.

According to one embodiment, the inner spar 462 may have a thickness approximately that of other internal structures. In particular, the inner spar 462 may have a wall thickness plus or minus 20% that of the one or more section dividers 464, one or more ribs 465. In addition, the inner spar 462 may be kept with 1.2 times the wall thickness of the skin 460.

According to one embodiment, the inner spar 462 may include one or more inner spar pass-through hole(s) 473. In particular, the inner spar 462 may include perforations such that pressure is equalized between the pressure side 448 (FIG. 5) and the lift side 449 (FIG. 5) of the inner spar 462. For example, an inner spar pass-through hole 473 may be made in each discrete sub-passageway or “section” of the single-bend heat exchange path 470. In addition, depending on the pressure profile of the particular cooled turbine blade 440, a single section may include more than one inner spar pass-through hole(s) 473. Furthermore, the inner spar pass-through hole(s) 473 may be located throughout the inner spar 462. For example, and as illustrated, the inner spar 462 may include inner spar pass-through hole(s) 473 near the platform 443, near the tip wall 461, and/or near the single bend.

Within the airfoil 441, each section divider 464 may extend from the base 442 to the trailing edge 447, generally including a ninety degree turn and including a smooth transition. In addition, each section divider 464 may extend outward from the inner spar 462 to the skin 460 on each of the pressure side 448 (FIG. 3) or the lift side 449 (FIG. 3). Accordingly, cooling air 15 may be constrained within a sub-passageway or “section” of the single-bend heat exchange path 470 defined by the inner spar 462, either of the pressure side 448 (FIG. 3) or the lift side 449 (FIG. 3) of the skin 460, a section divider 464, and one of: an adjacent section divider 464, the tip wall 461, and the base 442.

According to one embodiment, each section divider 464 on one side of inner spar 462 may run parallel with each other. According to another embodiment, a section divider 464 on the pressure side 448 (FIG. 3) of the inner spar 462 may minor another section divider 464 on the lift side 449 (FIG. 3) of the inner spar 462. Furthermore two “mirrored” section dividers 464 may merge into a single section divider 464 downstream of the inner spar 462 such that the “merged” section divider 464 extends from the pressure side 448 (FIG. 3) of the skin 460 directly to the lift side 449 (FIG. 3) of the skin 460.

Within the airfoil 441, each rib 465 may extend radially from the base 442 toward the tip end 445, terminating prior to reaching the tip wall 461. In addition, each rib 465 may extend outward from the inner spar 462 to the skin 460 on either of the pressure side 448 (FIG. 3) or the lift side 449 (FIG. 3) (i.e., in and out of plane). According to one embodiment, a rib 465 may also include a single bend at its distal end, relative to the base 442. The single bend may be approximately ninety degrees and include a smooth transition. In addition, the rib 465 may run parallel with an adjacent structure (e.g., section divider 464). Furthermore, and as above, a rib 465 on the pressure side 448 (FIG. 3) of the inner spar 462 may mirror another rib 465 on the lift side 449 (FIG. 3) of the inner spar 462.

According to one embodiment, the airfoil 441 may include a leading edge rib 472. The leading edge rib 472 may extend radially from the base 442 toward the tip end 445, terminating prior to reaching the tip wall 461. In addition, the leading edge rib 472 may extend directly from the pressure side 448 (FIG. 3) of the skin 460 to the lift side 449 (FIG. 3) of the skin 460. In doing so, the leading edge rib 472 may form the leading edge chamber 463 in conjunction with the skin 460 at the leading edge 446 of the airfoil 441. Additionally, all or part of the cooling air 15 leaving the leading edge chamber 463 may be redirected toward the trailing edge 447 by tip wall 461 and other cooling air 15 within the airfoil 441. Accordingly, the leading edge chamber 463 may form part of the single-bend heat exchange path 470.

Within the airfoil 441, each air deflector 466 may extend outward from the inner spar 462 to the skin 460 on either of the pressure side 448 (FIG. 3) or the lift side 449 (FIG. 3). Each air deflector 466 may include a single bend, which is configured to redirect cooling air 15 approximately ninety degrees. Accordingly, the single bend may be approximately ninety degrees and include a smooth transition. Generally, the single bend of the air deflector 466 may start from a radial/vertical direction and smoothly transition to a horizontal direction aimed toward the trailing edge 447. In addition, the single bend of the air deflector 466 may run parallel with the single bend of an adjacent section divider 464 or rib 465. Furthermore, and as above, an air deflector 466 on the pressure side 448 (FIG. 3) of the inner spar 462 may mirror another air deflector 466 on the lift side 449 (FIG. 3) of the inner spar 462.

According to one embodiment, the airfoil 441 may include a leading edge air deflector 475. As above, the leading edge air deflector 475 may include a single bend, which is configured to redirect cooling air 15 approximately ninety degrees. Accordingly, the single bend may be approximately ninety degrees and include a smooth transition. The leading edge air deflector 475 may be located so as to redirect cooling air 15 leaving the leading edge chamber 463. In particular, the leading edge air deflector 475 may be radially located between and the leading edge rib 472 and the tip wall 461. Additionally, the leading edge air deflector 475 may physically interact with the inner spar 462. In particular, the leading edge air deflector 475 may extend from the pressure side 448 (FIG. 3) of the skin 460 to the lift side 449 (FIG. 3) of the skin 460, wherein at least a portion of the leading edge air deflector 475 is intersected by the inner spar 462 between the pressure side 448 (FIG. 3) of the skin 460 and the lift side 449 (FIG. 3) of the skin 460.

Within the airfoil 441, the plurality of inner spar cooling fins 467 may extend outward from the inner spar 462 to the skin 460 on either of the pressure side 448 (FIG. 3) or the lift side 449 (FIG. 3). In contrast, the plurality of trailing edge cooling fins 469 may extend from the pressure side 448 (FIG. 3) of the skin 460 directly to the lift side 449 (FIG. 3) of the skin 460. Accordingly, the plurality of inner spar cooling fins 467 are located forward of the plurality of trailing edge cooling fins 469, as measured along the mean camber line 474 (FIG. 3) of the airfoil 441.

Both the inner spar cooling fins 467 and the trailing edge cooling fins 469 may be disbursed copiously throughout the single-bend heat exchange path 470. In particular, the inner spar cooling fins 467 and the trailing edge cooling fins 469 may be disbursed throughout the airfoil 441 so as to thermally interact with the cooling air 15 for increased cooling. In addition, the distribution may be in the radial direction and in the direction along the mean camber line 474 (FIG. 3). The distribution may be regular, irregular, staggered, and/or localized.

According to one embodiment, the inner spar cooling fins 467 may be long and thin. In particular, inner spar cooling fins 467, traversing less than half the thickness of the airfoil 467, may use a round “pin” fin. Moreover, pin fins having a height-to-diameter ratio of 2-7 may be used. For example, the inner spar cooling fins 467 may be pin fins having a diameter of 0.017-0.040 inches, and a length off the inner spar 467 of 0.034-0.240 inches.

According to one embodiment, the inner spar cooling fins 467 may also be densely packed. In particular, inner spar cooling fins 467 may be within two diameters of each other. Thus, a greater number of inner spar cooling fins 467 may be used for increased cooling. For example, across the inner spar 462, the fin density may be in the range of 80 to 300 fins per square inch per side of the inner spar 462. Alternately, the fin density may be in the range of 50 to 400 fins per square inch per side of the inner spar 462. Alternately, the fin density may be in the range of 40 to 70 fins per square inch per side of the inner spar 462. Alternately, the fin density may be in the range of 215 to 385 fins per square inch per side of the inner spar 462. Alternately, the fin density may be in the range of 70 to 385 fins per square inch per side of the inner spar 462. Alternately, the fin density may be in the range of 40 to 215 fins per square inch per side of the inner spar 462.

Within the airfoil 441, the trailing edge rib 468 may extend radially from the base 442 toward the tip end 445. In particular, the trailing edge rib 468 may radially extend between the base 442 and the section divider 464 that defines the subdivision of the single-bend heat exchange path that exhausts nearest the platform 443. In addition, the trailing edge rib 468 may be located along the inner spar trailing edge 476 and between the inner spar cooling fins 467 and the trailing edge cooling fins 469.

Unlike a section divider 464 or a rib 465, the trailing edge rib 468 may be perforated to include one or more openings. This will allow cooling air 15 to pass through the trailing edge rib 468 toward the cooling air outlet 471 in the trailing edge 447, and thus complete the single-bend heat exchange path 470.

Taken as a whole the cooling air passageway 482 and the single-bend heat exchange path 470 may be coordinated. In particular and returning to the base 442 of the cooled turbine blade 440, the cooling air passageway 482 may be sub-divided into a plurality of flow paths. As illustrated, the subdivided cooling air passageway 482 may be coordinated with the one or more section divider(s) 464 and the one or more rib(s) 465 above, in the airfoil 441. Accordingly, each subdivision within the base 442 may be aligned with and include a cross sectional shape (not shown) corresponding to the areas bounded by the skin 460 and each section divider 464 and rib 465. In addition, the cooling air passageway 482 may maintain the same overall cross sectional area (i.e., constant flow rate and pressure) in each subdivision, as between the cooling air inlet 481 and the airfoil 441. Alternately, the cooling air passageway 482 may vary the cross sectional area of individual subdivisions where differing performance parameters are desired for each section, in a particular application.

According to one embodiment, the cooling air passageway 482 and the single-bend heat exchange path 470 may each include asymmetric divisions for reflecting localized thermodynamic flow performance requirements. In particular, as illustrated and discussed above, the cooled turbine blade 440 may have two or more sections divided by the one or more section divider(s) 464. Accordingly, there will be a section on each side of the section divider 464. As with the cooling air passageway 482, each section may maintain the same overall cross sectional area. Alternately, each section divider 464 may be located such that each section varies where different performance parameters are desired for each section, in a particular application. For example, by moving the horizontal arm of section divider 464 radially outward, and a larger section is created on its inward side, and vis versa.

Similarly, according one embodiment, the individual inner spar cooling fins 467 and the trailing edge cooling fins 469 may also include localized thermodynamic structural variations. In particular, the inner spar cooling fins 467 and/or the trailing edge cooling fins 469 may have different cross sections/surface area and/or fin spacing at different locations of the inner spar 462. For example, the cooled turbine blade 440 may have localized “hot spots” that favor a greater thermal conductivity, or low internal flow areas that favor reduced airflow resistance. In which case, the individual cooling fins may be modified in shape, size, positioning, spacing, and grouping.

According to one embodiment, one or more of the inner spar cooling fins 467 and the trailing edge cooling fins 469 may be pin fins or pedestals. The pin fins or pedestals may include many different cross-sectional areas, such as: circular, oval, racetrack, square, rectangular, diamond cross-sections, just to mention only a few. As discussed above, the pin fins or pedestals may be arranged as a staggered array, a linear array, or an irregular array.

FIG. 5 is a sectional top view of the turbine blade of FIG. 4, as taken along plane indicated by broken line 5-5 of FIG. 4. From this view, inner spar 462 and the relationship with the above features and structures within the airfoil 441 are shown. For clarity, only the nearest row of internal structures within the airfoil 441 is shown. In addition, some of the cutaway internal structures are illustrated with alternating hatching for convenience and clarity, however, as discussed herein, in different embodiments they may be made from the same or different materials.

As illustrated, airfoil 441 may have a varying profile in the radial direction. In particular, airfoil 441 may have a greater thickness near the platform 443 of base 442 than near the tip end 445 (FIG. 3), as can be seen viewing both FIG. 3 (showing the airfoil 441 at the tip end 445) and FIG. 5 (showing the airfoil 441 closer to the base 442). The illustrated shape of the airfoil 441 is merely representative, and may vary from application to application. Moreover, airfoil 441 may retain its aerodynamic features (i.e., leading edge 446, trailing edge 447, pressure side 448, lift side 449, and mean camber line 474) independent of its particular shape. Also, the illustrated thickness of the skin 460 and the structures residing within are also representative and not limiting.

As illustrated, inner spar 462 may be located in between the pressure side 448 of the skin 460 and the lift side 449 the skin 460. In particular, the inner spar 462 may substantially coincide with the mean camber line 474 of the airfoil 441. Accordingly, inner spar 462 may bifurcate the single-bend heat exchange path 470 into a cavity associated with the pressure side 448 of the airfoil 441 and a cavity associated with the lift side 449 of the airfoil 441. Moreover, each section divider 464 and each rib 465 may further sub-divide the single-bend heat exchange path 470. In particular and as discussed above, each section divider 464 and each rib 465 may extend outward from the inner spar 462 to the skin 460 on both the pressure side 448 and the lift side 449, limiting cross flow within the single-bend heat exchange path 470 and subdividing the cavity on the pressure side 448 on the lift side 449 into a series of generally parallel cavities/flow passages.

According to one embodiment, inner spar 462 may extend between the leading edge chamber 463, at the leading edge rib 472, and the trailing edge rib 468. As above and as illustrated, leading edge rib 472 and the trailing edge rib 468 may each extend from the pressure side 448 of the skin 460 directly to the lift side 449 of the skin 460. Accordingly, the forward and aft ends of the inner spar 462 may be bound along the mean camber line 474 by the leading edge rib 472 and the trailing edge rib 468, respectively. Notably, the origination of the inner spar 462 at the leading edge rib 472 provides for an increased cross section of the leading edge chamber 463. Notwithstanding, according to one embodiment, the inner spar 462 may extend at least seventy-five percent the length of the mean camber line 474.

As illustrated and discussed above, inner spar 462 may support the extension of the one or more section dividers 464, the one or more ribs 465, the one or more air deflectors 466, and the plurality of inner spar cooling fins 467. In particular, each structure/feature may extend from the inner spar 462 to the pressure side 448 or the lift side 449 of the airfoil 441. According to another embodiment, each structure/feature may run parallel to each other. Likewise, each structure/feature may be oriented perpendicular to the forward edge 484 (of aft edge 485) of the platform 443, which may also be viewed as perpendicular to the center axis 95 (FIG. 1).

For convenience or clarity, and as the entire cooled turbine blade 440 may be formed as a single casting, each structure/feature having a mirror structure/feature opposite the inner spar 462 may be equally treated or referred to as a single member or as two separate members. For example, section dividers 464 on both sides of the inner spar 462 may equally be described as two separated members (i.e., as a first section divider 464 extending from the inner spar 462 to the lift side 449 of the skin 460 and a second section divider 464 extending from the inner spar 462 to the pressure side 449 of the skin 460) or as a single member that passes through or includes the corresponding section of the inner spar 462 (i.e., as a section divider 464 extending between the skin 460 on the lift side 449 and to the skin 460 on the pressure side 448).

According to one embodiment and as illustrated each structure/feature may include a “mirror image” on the opposite side of the inner spar 462. Notably, as the section cut is taken radially inward of the single bend of the section dividers 464, only a portion is illustrated. As discussed above each section divider 464 may extend to the trailing edge 447, and two “mirrored” section dividers 464 may merge into a single section divider 464 downstream of the inner spar 462 such that the “merged” section divider 464 extends from the pressure side 448 of the skin 460 directly to the lift side 449 of the skin 460.

Both the inner spar cooling fins 467 and the trailing edge cooling fins 469 may be oriented for thermal performance, structural performance, and/or manufacturability. For example, the plurality of inner spar cooling fins 467 may be oriented substantially parallel to each other and perpendicular to the center axis 95. In addition, plurality of inner spar cooling fins 467 may populate at least ten percent of the volume of the single-bend heat exchange path 470. Also, the plurality of first inner spar cooling fins 467 may have a length at least twenty-five percent longer than the thickness of the inner spar 462, as measured between the inner spar 462 and the pressure side 448 or the lift side 449 of the airfoil 441.

With regard to the structures/features toward the trailing edge 447 of the airfoil 441, having a narrower thickness, the structures/features may extend directly from the pressure side 448 to the lift side 449 of the skin 460. In particular, both the trailing edge rib 468 and the plurality of trailing edge cooling fins 469 may extend skin-to-skin. Like the inner spar cooling fins 467, the plurality of trailing edge cooling fins 469 may be oriented substantially parallel to each other. However, trailing edge cooling fins 469 may also be oriented so as to reduce the distance of the span between the pressure side 448 and the lift side 449 of the skin 460. For example, the plurality of trailing edge cooling fins 469 may be oriented substantially perpendicular to the mean camber line 474. Alternately, the plurality of trailing edge cooling fins 469 may be oriented substantially perpendicular to the skin 460 of the airfoil 441 as averaged between the pressure side 448 and the lift side 449.

INDUSTRIAL APPLICABILITY

The present disclosure generally applies to cooled turbine blades, and gas turbine engines having cooled turbine blades. The described embodiments are not limited to use in conjunction with a particular type of gas turbine engine, but rather may be applied to stationary or motive gas turbine engines, or any variant thereof. Gas turbine engines, and thus their components, may be suited for any number of industrial applications, such as, but not limited to, various aspects of the oil and natural gas industry (including include transmission, gathering, storage, withdrawal, and lifting of oil and natural gas), power generation industry, cogeneration, aerospace and transportation industry, to name a few examples.

Generally, embodiments of the presently disclosed cooled turbine blades are applicable to the use, assembly, manufacture, operation, maintenance, repair, and improvement of gas turbine engines, and may be used in order to improve performance and efficiency, decrease maintenance and repair, and/or lower costs. In addition, embodiments of the presently disclosed cooled turbine blades may be applicable at any stage of the gas turbine engine's life, from design to prototyping and first manufacture, and onward to end of life. Accordingly, the cooled turbine blades may be used in a first product, as a retrofit or enhancement to existing gas turbine engine, as a preventative measure, or even in response to an event. This is particularly true as the presently disclosed cooled turbine blades may conveniently include identical interfaces to be interchangeable with an earlier type of cooled turbine blades.

As discussed above, the entire cooled turbine blade may be cast formed. According to one embodiment, the cooled turbine blade 440 may be made from an investment casting process. For example, the entire cooled turbine blade 440 may be cast from stainless steel and/or a superalloy using a ceramic core or fugitive pattern. Accordingly, the inclusion of the inner spar is amenable to the manufacturing process. Notably, while the structures/features have been described above as discrete members for clarity, as a single casting, the structures/features may pass through and be integrated with the inner spar. Alternately, certain structures/features (e.g., skin 460) may be added to a cast core, forming a composite structure.

Embodiments of the presently disclosed cooled turbine blades provide for a lower pressure cooling air supply, which makes it more amenable to stationary gas turbine engine applications. In particular, the single bend provides for less turning losses, compared to serpentine configurations. In addition, the inner spar and copious cooling fin population provides for substantial heat exchange during the single pass. In addition, besides structurally supporting the cooling fins, the inner spar itself may serve as a heat exchanger. Finally, by including subdivided sections of both the single-bend heat exchange path in the airfoil, and the cooling air passageway in the base, the cooled turbine blades may be tunable so as to be responsive to local hot spots or cooling needs at design, or empirically discovered, post-production.

The disclosed single-bend heat exchange path 470 begins at the base 442 where pressurized cooling air 15 is received into the airfoil 441. The cooling air 15 is received from the cooling air passageway 482 in a generally radial direction. The single-bend heat exchange path 470 is configured such that cooling air 15 will pass between, along, and around the various internal structures, but will generally flow in a ninety degree path as viewed from the side view (conceptually treating the camber sheet as a plane). Accordingly, the single-bend heat exchange path 470 may include some negligible lateral travel (i.e., into the plane) associated with the general curvature of the airfoil 441. Also, as discussed above, although the single-bend heat exchange path 470 is illustrated by a single representative flow line traveling through a single section for clarity, the single-bend heat exchange path 470 includes the entire flow path carrying cooling air 15 through the airfoil 441. Moreover, unlike other internally cooled turbine blades, the single-bend heat exchange path 470 is not serpentine, but rather has a single bend that efficiently redirects the cooling air 15 to the cooling air outlet 471 at the trailing edge 447 with a single turn.

Although this invention has been shown and described with respect to detailed embodiments thereof, it will be understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the spirit and scope of the claimed invention. Accordingly, the preceding detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. In particular, the described embodiments are not limited to use in conjunction with a particular type of gas turbine engine. For example, the described embodiments may be applied to stationary or motive gas turbine engines, or any variant thereof. Furthermore, there is no intention to be bound by any theory presented in any preceding section. It is also understood that the illustrations may include exaggerated dimensions and graphical representation to better illustrate the referenced items shown, and are not consider limiting unless expressly stated as such. 

What is claimed is:
 1. A turbine blade for use in a gas turbine engine, the turbine blade comprising: a base; an airfoil comprising a skin extending from the base and forming a leading edge, a trailing edge, a pressure side, and a lift side; a plurality of trailing edge cooling fins extending from the pressure side of the skin to the lift side of the skin; a leading edge rib, the leading edge rib extending from the pressure side of the skin to the lift side of the skin, the leading edge rib extending from the base, proximal and spaced apart from the leading edge and within the skin; an inner spar within the skin, the inner spar extending from the leading edge rib toward the trailing edge; a plurality of first inner spar cooling fins extending from the inner spar to the skin on the pressure side of the airfoil; and a plurality of second inner spar cooling fins extending from the inner spar to the skin on the lift side of the airfoil.
 2. The turbine blade of claim 1, wherein the inner spar extends from the leading edge rib toward the trailing edge substantially along a mean camber line of the airfoil to include a length of at least seventy-five percent of the mean camber line length between the leading edge and the trailing edge, and terminating prior to reaching the plurality of trailing edge cooling fins.
 3. The turbine blade of claim 1, wherein the plurality of first inner spar cooling fins and the plurality of second inner spar cooling fins are oriented substantially parallel to each other.
 4. The turbine blade of claim 1, wherein the base includes at least one cooling air passageway, the turbine blade further comprising: a single-bend heat exchange path within the airfoil, the single-bend heat exchange path interfacing with and beginning at the at least one cooling air passageway in the base, and terminating at the trailing edge, the single-bend heat exchange path configured to redirect the cooling air from the at least one cooling air passageway in the base toward the trailing edge; and wherein the single-bend heat exchange path is further configured to redirect the cooling air such that the cooling air is redirected in a single turn; and wherein at least a portion of the single-bend heat exchange path is sub-divided by the inner spar.
 5. The turbine blade of claim 1, wherein the plurality of first inner spar cooling fins extend from the inner spar with a density of at least 80 fins per square inch; and wherein the plurality of second inner spar cooling fins extend from the inner spar with a density of at least 80 fins per square inch.
 6. The turbine blade of claim 1, wherein the turbine blade is cast from a single material; and wherein the plurality of first inner spar cooling fins and the plurality of second inner spar cooling fins have a length at least twenty-five percent longer than the thickness of the inner spar.
 7. A gas turbine engine including the turbine blade of claim
 1. 8. A turbine blade for use in a gas turbine engine, the turbine blade comprising: a base; an airfoil comprising a skin extending from the base and forming a leading edge, a trailing edge, a pressure side, a lift side, and a tip end, the tip end located distally from the base; a plurality of trailing edge cooling fins extending from the pressure side of the skin to the lift side of the skin; an inner spar extending from the base toward the tip end, the inner spar located between the pressure side of the skin and the lift side the skin; a plurality of first inner spar cooling fins extending from the inner spar to the skin on the pressure side of the airfoil, the plurality of first inner spar cooling fins oriented substantially parallel to each other; and a plurality of second inner spar cooling fins extending from the inner spar to the skin on the lift side of the airfoil, the plurality of second inner spar cooling fins oriented substantially parallel to each other and with the plurality of first inner spar cooling fins.
 9. The turbine blade of claim 8, wherein the base includes a platform having a forward edge; and wherein the plurality of first inner spar cooling fins and the plurality of second inner spar cooling fins are oriented substantially perpendicular to the forward edge.
 10. The turbine blade of claim 8, further comprising: a single-bend heat exchange path within the airfoil, the single-bend heat exchange path interfacing with and beginning at an at least one cooling air passageway in the base, and terminating at the trailing edge, the single-bend heat exchange path configured to redirect the cooling air from the at least one cooling air passageway in the base toward the trailing edge substantially along a mean camber line; and wherein the single-bend heat exchange path is further configured to redirect the cooling air such that the cooling air is redirected in a single turn; and wherein at least a portion of the single-bend heat exchange path is sub-divided by the inner spar.
 11. The turbine blade of claim 8, wherein the plurality of first inner spar cooling fins extend from the inner spar with a density of at least 80 fins per square inch; and wherein the plurality of second inner spar cooling fins extend from the inner spar with a density of at least 80 fins per square inch.
 12. The turbine blade of claim 12, wherein the turbine blade is cast from a single material; and wherein the plurality of first inner spar cooling fins and the plurality of second inner spar cooling fins have a length at least twenty-five percent longer than the thickness of the inner spar.
 13. A gas turbine engine including a turbine having a turbine rotor assembly that includes the turbine blade of claim 1; and wherein the turbine rotor assembly is installed in a first stage of the turbine.
 14. A turbine blade for use in a gas turbine engine, the turbine blade comprising: a base; an airfoil comprising a skin extending from the base and forming a leading edge, a trailing edge, a pressure side, a lift side, and a tip end distal from the base; a plurality of trailing edge cooling fins extending from the pressure side of the skin to the lift side of the skin; an inner spar extending from the base toward the tip end, the inner spar located between the pressure side of the skin and the lift side the skin; a plurality of first inner spar cooling fins extending from the inner spar to the skin on the pressure side of the airfoil; a plurality of second inner spar cooling fins extending from the inner spar to the skin on the lift side of the airfoil; and a single-bend heat exchange path within the airfoil, the single-bend heat exchange path interfacing with and beginning at the at least one cooling air passageway in the base, and terminating at the trailing edge, at least a portion of the single-bend heat exchange path being sub-divided by the inner spar, the single-bend heat exchange path configured to redirect the cooling air from the at least one cooling air passageway in the base toward the trailing edge; the single-bend heat exchange path further configured to redirect the cooling air such that the cooling air is redirected in a single turn.
 15. The turbine blade of claim 14, further comprising: at least one first section divider extending from the base to the trailing edge while substantially following a ninety degree path, the at least one first section divider further extending from the inner spar to the skin on pressure side of the airfoil; and at least one second section divider extending from the base to the trailing edge while substantially following a ninety degree path, the at least one first section divider further extending from the inner spar to the skin on the lift side of the airfoil; and wherein the at least one first section divider and the at least one second section divider are joined together downstream of the inner spar to form a single section divider extending from the pressure side of the skin directly to the lift side of the skin.
 16. The turbine blade of claim 15, further comprising: at least one first rib radially extending from the base; the at least one first rib further extending from the inner spar to the skin on pressure side of the airfoil; and at least one second rib radially extending from the base; the at least one second rib further extending from the inner spar to the skin on the lift side of the airfoil.
 17. The turbine blade of claim 16, further comprising: at least one first air deflector extending from the inner spar to the pressure side of the airfoil, the at least one first air deflector located within the single-bend heat exchange path and configured to redirect cooling air toward the trailing edge; and at least one second air deflector extending from the inner spar to the lift side of the airfoil, the at least one first air deflector located within the single-bend heat exchange path and configured to redirect cooling air toward the trailing edge.
 18. The turbine blade of claim 14, wherein the plurality of first inner spar cooling fins extend from the inner spar with a density of at least 80 fins per square inch; and wherein the plurality of second inner spar cooling fins extend from the inner spar with a density of at least 80 fins per square inch.
 19. The turbine blade of claim 14, wherein the turbine blade is cast from a single material.
 20. The turbine blade of claim 19, wherein the plurality of first inner spar cooling fins and the plurality of second inner spar cooling fins have a length at least twenty-five percent longer than the thickness of the inner spar. 